The present invention relates to a mechanism making it possible to automatically open out or extend an appendage, such as a solar panel, on a space vehicle, more particularly constituted by a satellite performing a rotary movement on itself.
The placing into orbit of an artificial satellite about the earth takes place in several stages. Thus, in the case of a geostationary satellite, the latter is generally placed on a low quasi-circular orbit at the time of its launch and the altitude thereof is e.g. approximately 200 km. It is only subsequently transferred to its final orbit, which is in this case constituted by a quasi-circular geostationary orbit at approximately 36,000 km from the earth.
The passage from the low orbit to the geostationary orbit takes place during a so-called transfer phase. The latter firstly comprises so-called perigee operations, during which the satellite passes from the low to an elliptical transfer orbit, whose perigee is located on the geostationary orbit to be reached. The passage from the elliptical transfer orbit to the geostationary orbit takes place during apogee operations constituting the second part of the transfer phase.
During this transfer phase and particularly during the apogee operations, the satellite is subject to high mechanical stressing. However, the existing mechanisms for extending solar panels generally incorporate control means and structures which it would be too constraining to dimension in order to withstand such mechanical stresses if extension took place during the transfer phase. Therefore these mechanisms are only intended to operate following the placing of the satellite in the geostationary orbit.
During the transfer phase, the operational satellite equipment is generally in the standby state or is out of operation. However, certain vital members for the survival and monitoring of the satellite requiring electrical energy must be able to operate. It is therefore necessary to have a certain electrical energy or power quantity during this transfer phase.
Moreover, the transfer phase can last a relatively long time, because it can involve the passage through several elliptical transfer orbits, whose unitary duration is approximately 10 hours. Thus, the electric power cannot exclusively be provided by secondary electricity sources, such as the chemical batteries equipping the satellite, because this would lead to excessive weight and overall dimensions.
In the case of an artificial satellite stabilized according to three axes during the transfer phase, French patent application No. 2 505 288 of the present Applicant proposes solving this problem by effecting the extension of the solar panels as soon as the perigee operation has been completed.
However, in certain cases of satellites spinning during the transfer phase, i.e. a satellite performing a rotary movement about its own axis during this phase and which can e.g. be the longitudinal axis of the satellite, the solution proposed by the aforementioned specification is unsatisfactory. Thus, the electrical energy supplied by the opening out of a single panel of each of the wing or fin members may then be inadequate. In this case, it is consequently necessary to increase the surface of the solar cells exposed during the transfer phase.